Rocket engine nozzle comprising a jet separation control system

ABSTRACT

The invention relates to a rocket engine nozzle comprising a system for controlling jet separation of the flow in the nozzle, wherein said control system exhibits a plurality of separation triggering elements ( 5, 10 ) arranged in such a way as to generate, from mutually spaced initiation points ( 9 ), distinct zones ( 6 ) of jet separation, so as to form a three-dimensional separation of the flow.  
     The flow control system can exhibit at least two triggering elements ( 5, 10 ).

[0001] The subject of the present invention is a rocket engine nozzle,exhibiting a jet separation control system, for example a device forinjecting fluid through a wall of the nozzle, so as to induce jetseparation in the gases ejected by the nozzle.

[0002] An important point in the design of a launcher is theoptimization of the performance of its engines. In particular, thenozzle must be designed so as to yield a maximum thrust coefficientcompatible with the limits imposed by the other constraints.

[0003] The thrust coefficient C of a nozzle is an increasing function ofthe ratio of the exit area Ae of the nozzle to the area At of the throatof the nozzle.

[0004] For an upper stage, which is ignited outside the atmosphere, thestatic pressure of the jet at the exit of the nozzle may be very low.The area ratio R=Ae/At of the nozzle is in this case essentially limitedby the space available.

[0005] On the other hand, when the nozzle operates within theatmosphere, the gases which exit the nozzle cannot expand to below alimit pressure Psep, at which a separation of flow in the nozzle occursspontaneously.

[0006] This jet separation is naturally unstable and generatesconsiderable aerodynamic forces at the moment of ignition and during theinitial atmospheric flight, which may even lead to the destruction ofthe nozzle if the jet separation is too considerable.

[0007] As far as engines which are designed to operate right from theground and to accomplish the major part of their mission outside theatmosphere are concerned, the determination of the ratio Ae/Atrepresents a difficult compromise.

[0008] Numerous devices have been proposed for controlling jetseparation in nozzles.

[0009] An up-to-date review of this topic has appeared in particular inthe article entitled Advanced Rocket Nozzles by Gerald Hagemann et al.,published in the Journal of Propulsion and Power, vol. 14 No. 5,September-October 1998, pages 620 to 634.

[0010] This deals in particular with “dual-bell” nozzles, nozzles withfixed or temporary inserts, two-position or extendible nozzles, externalexpansion nozzles, so-called expansion/deflection nozzles, nozzlesexhibiting a variable throat area, and finally dual-mode nozzles.

[0011] The control of jet separation in a nozzle with the aid ofsecondary injection of gas has also been proposed, but this secondaryinjection has the effect of preserving axial symmetry of the flow. Thistechnique is recalled in point 4, page 626 of the aforesaid article.

[0012] Experiments carried out on an RL10 engine and implementingpassive injection are described in the article entitled “AltitudeCompensating Nozzle Evaluation” by R. C. PARSLEY et al., published inthe proceedings of the 28^(th) Joint Propulsion Conference and Exhibit,6 to Jul. 8, 1992, Nashville, Tenn., pages 1 to 6.

[0013] Finally, American patent U.S. Pat. No. 3,925,982 (Martin MariettaCorporation) describes a rocket engine exhibiting a high nozzle arearatio and which is equipped with a device for active secondary injectionexhibiting a shock generating ring which is intended to control jetseparation, by forcing the boundary layer of the primary gas jet toseparate uniformly from the wall of the nozzle.

[0014] This is achieved with the aid of a large number of injectionpoints which are distributed around the circumference of the nozzle.These injection points are closely spaced, and they inject a secondarygas jet radially and inwardly of the nozzle so as to effect jetseparation which is invariant with any rotation about the axis of thenozzle.

[0015] Alternatively, this jet separation can be achieved via acontinuous slot extending over the entire circumference of the nozzle.

[0016] The theory of jet separation has been recalled in the recentarticle by G. L. ROMINE entitled “Nozzle Flow Separation” published inthe AIAA Journal, vol. 36, No. 9, September 1998, pp. 1618-1625.

[0017] The theory of secondary injection has been set out in the articleentitled “Some aspects of gaseous secondary injection with applicationto thrust vector control” by R. D. GUHSE et al., published inproceedings No. 71-750 of the AIAA/SAE 7^(th) Propulsion JointSpecialist Conference of Salt Lake City, Jun. 14-18, 1971, pages 1 to 8.

[0018] The known techniques of secondary injection, which involve jetseparation exhibiting axial symmetry, that is to say which is invariantabout any rotation about the axis of the nozzle, exhibit the followingproblems:

[0019] active secondary injection is difficult to implement, given thatthe mass flux which is required for effective generation of axialsymmetric jet separation is high;

[0020] passive secondary injection which implements ventilation of thenozzle is operational only within a limited range of differentialpressure, which implies that in order to obtain a nozzle which operatesat all altitudes, its porosity must be continuously variable as afunction of external pressure and of the operational parameters of theengine, this being hardly compatible with the nozzle constructionconstraints.

[0021] One of the drawbacks of secondary injections with axial symmetry,such as for example that described in the aforesaid American patent, isthat under certain engine operating conditions, the jet separationcommences at a random point on the injection ring, and whose position,which depends on the upstream disturbances, is unstable.

[0022] An object of the invention is to propose a separation controlsystem, especially through secondary injection which avoids suchinstability.

[0023] Another object of the invention is to reduce the unsteady loadsapplied to nozzles and hence to reduce the necessary mechanical strengthof the nozzles and the engine mountings, thereby making it possible toreduce their mass.

[0024] Another object of the invention is to allow the installation onrocket engines used right from the ground, of nozzles with a higherexpansion ratio and hence to enable an overall improvement in theperformance of these engines.

[0025] Another object of the invention is to minimize the totalsecondary injection flux required to obtain stable separation.

[0026] Another object of the invention is to control the separation ofthe jet during ignition of the engine on the ground.

[0027] Another object of the invention is to be able to facilitate thecontrol of jet separation as a function of altitude.

[0028] At least one of the aforesaid objects of the present invention isachieved through a rocket engine nozzle comprising a system forcontrolling jet separation, wherein said control system exhibits aplurality of separation triggering elements arranged in such a way as togenerate, from mutually spaced initiation points, distinct zones of jetseparation, so as to form a three-dimensional separation of the flow.

[0029] According to a first variant, the invention relates to anejection nozzle for a rocket engine exhibiting a device for injectingfluid through a wall of the nozzle, so as to induce jet separation inthe gases ejected by the nozzle, in which the control system is afluid-injection device which exhibits in at least one injection crosssection perpendicular to the axis of the nozzle, at least twoindependent injection orifices distributed over the perimeter of thewall of the nozzle, each injection orifice constituting a saidseparation triggering element inducing a said distinct zone of jetseparation.

[0030] According to the invention, each separation triggering element,for example each injection orifice initiates and maintains locally theseparation of the jet, thereby remedying the aforesaid instability. Thisarrangement is essentially different from that which is described in theaforesaid American patent, for which the injection orifices are tightlydistributed over the perimeter of the nozzle so as to generate a flowseparation which is invariant with any rotation of the axis of thenozzle, and which operates in a manner equivalent to that of a ringexhibiting a continuous slot extending over the entire circumference ofthe nozzle.

[0031] It is advantageous for the injection orifices, for example two innumber or else three in number, to be uniformly distributed over theperimeter of the wall of the nozzle. This makes it possible to avoid toa large extent the occurrence of transverse forces applied to thenozzle.

[0032] Said injection cross section is advantageously arranged at adistance D from the throat of the nozzle which is substantially lessthan the distance Do of spontaneous separation of the jet.

[0033] Said nozzle cross section is preferably chosen at a level wherethe static pressure P of the jet is substantially greater than thenatural separation pressure of the jet Psep, for example P=2 Psep.

[0034] An injection device can exhibit a plurality of injection crosssections situated at different distances from the throat, and adistributing device for feeding one or other of the injection crosssections, in such a way as to take into account, in a manner known perse, the variation as a function of altitude of the cross section where aso-called spontaneous separation of the jet occurs.

[0035] According to an embodiment making it possible to control jetseparation during ignition of the engine on the ground, the flow controlsystem exhibits an external stabilizing device integral with aground-based installation and which exhibits, on the one hand, a numberN(N≧2) of injection tubes (for example parallel to the axis of thenozzle) each of which constitutes a said separation triggering element,and which are distributed, preferably downstream of the nozzle, in sucha way as to direct in counter-current to the main stream of the nozzlestabilizing fluid jets toward N impact points situated downstream of thethroat of the nozzle, and on the other hand, a device for feeding theinjection tubes so as to feed them with fluid for a predeterminedtransient duration of ignition before takeoff, with a flow rate which issufficient for each impact point to induce a different zone of jetseparation of the nozzle.

[0036] The injection tubes are preferably arranged at the outlet of thenozzle exit.

[0037] The injection points of the injector of the external stabilizingdevice are preferably uniformly distributed over the perimeter of thewall of the nozzle. They are advantageously two in number (diametricallyopposed) or three in number (distributed at around 120° over theperimeter of the nozzle).

[0038] Other characteristics and advantages of the invention will becomemore apparent on reading the description which will follow, given by wayof non-limiting example in conjunction with the drawings hereinappended, in which:

[0039]FIG. 1 represents a device for implementing the present invention;and

[0040]FIG. 2 represents an additional device according to the presentinvention which is implemented during engine start-up.

[0041] As shown in FIG. 1, a nozzle, designated by the general label 1,exhibits a combustion chamber 2, a throat 3, and a divergent nozzle body4 which terminates in an exit cross section 8.

[0042] Over the perimeter of the divergent portion 4 of the nozzle, andin a cross section 7 situated in a plane, perpendicular to the axis ofthe nozzle, where the static pressure P of the jet is substantiallygreater than the nozzle separation pressure Psep, are arranged injectionorifices 5 able to direct radially inward a jet of a fluid, for examplethe combustion gases originating from the turbopumps of the engine.

[0043] The flow separation which is generated by these orifices 5, doesnot exhibit axial symmetry, but on the contrary it is three-dimensional.This is because each of the injection points 5, represented here asthree in number and distributed uniformly at 120° around the contour ofthe body 4 of the nozzle, induces a region of separation 6 of the streamexiting the nozzle. Owing to the determination of a limited number ofinjection points 5 which induce an equal number of separation regions 6,the position of the points of initial separation is not indeterminateand this makes it possible to solve the problem of instability.

[0044] Furthermore, by reason of the uniform distribution of theinjection points 5 around the circumference of the nozzle body 4 in theplane 7, the resultant of the lateral forces which are exerted on thenozzle and which, in the prior art is unstable, remains close to theaxis of the nozzle.

[0045] The cross section in which the injection is carried out is chosenin such a way as to be of slightly smaller area than that of the crosssection at which spontaneous stream separation would be apt to occur atlow altitude.

[0046] The device described makes it possible to obtainthree-dimensional separation exhibiting a plurality of separated regions6 which may possibly remerge downstream toward the exit 8 of the nozzle.

[0047] In theory, the number of injection points 5 could be equal tojust 2 so as to make it possible to maintain a symmetric thrust for thenozzle. The number of three injection points 5 seems however to be apreferable choice in order to avoid accidental separation of half thenozzle which could arise on ignition.

[0048] Moreover, a higher number than three injection points may beenvisaged, but this does not afford any appreciable advantage. In anyevent, the number and the spacing of the injection points must be chosenin such a way as to avoid any continuity of jet separation, which wouldamount in fact to the operating conditions equivalent to that of auniform ring (see the aforesaid American patent U.S. Pat. No.3,925,982).

[0049] The manner of operation of the invention can be explained asfollows:

[0050] The gases which are generated in the high-pressure combustionchamber 2 are, after passing through the throat 3, subjected to anexpansion in the ejection nozzle 4 and the static pressure decreases atthe same time as the jet is directed toward the exit 8 of the nozzle.

[0051] When a rocket engine operates at the static pressure whichprevails at sea level, and without implementing secondary injection, assoon as the static pressure of the jet approaches the normal pressurefor separation, the jet in the nozzle is apt to separate spontaneously,but such separation is unstable and occurs in a random place of a crosssection of the body of the nozzle 4, possibly creating appreciableunsteady forces.

[0052] The injection orifices 5 which direct a secondary stream radiallyinward according to the invention create an obstacle to the main jetlocally, thereby creating an arc-shaped shock wave in respect of theincident supersonic jet. This arc-shaped shock zone interacts with theboundary layer, in which it creates an increase in pressure justupstream of the injection point 5, thereby inducing local separation ofthe boundary layer at the points 9. Given that the boundary layer wasalready under conditions under which it was near to spontaneousseparation, the jet of the nozzle cannot reattach itself to the wall ofthe body 4 and the separation of the boundary layer spreads so as toadopt for each injection point 5 a conical configuration as shown by thedashed drawing of the separation regions 6 in FIG. 1. The vertex of thecones 6 is constituted by the points 9 of initiation of jet separation.The three injection points 5 create, starting from the initiation points9, three substantially identical cones 6 which are apt to remergedownstream so as to form an entirely separated jet at the exit 8 of thenozzle.

[0053] Given that the points of initiation 9 where the jet separationsoccur are imposed geometrically by the position of the three injectionpoints 5, the symmetry of revolution is broken and the points 9 ofinitiation of the separation are stable over time. The shocks which arecreated due to the separation of the boundary layer relative to the wallof the body 4 also remain localized and the residual vibrations due tothese shocks are of low amplitude, as are the residual unsteady forces.

[0054] Furthermore, the injected flux which is required for achievingseparation according to the invention is in principle smaller than inthe case of secondary injection with axial symmetry of the prior art.This is because such an injection of the prior art implements a largenumber of orifices along a circular cross section of the nozzle. Theseparation of the jet requires locally at each of these numerousorifices the same minimum flow rate as each of the few orifices usedaccording to the present invention. In the present invention,separation, which is achieved only onward of the localized points ofinitiation 9, tends thereafter to self-propagate along the periphery ofthe nozzle along with the widening and merging of the cones 6.

[0055] The process according to the invention is particularly beneficialin respect of engines exhibiting a gas generator. Such an engineexhibits turbopumps which discharge hot gases at a pressure greater thanatmospheric pressure. The Vulcain engine is of this type. The gases fromthe turbine of the Vulcain 1 engine are expelled from the engine. In theVulcain 2 engine, pipelines are already installed for reinjecting theturbine exit gases into the divergent portion 4 with a uniformdistribution along a cross section of the nozzle, so as to cool thedivergent portion, but nevertheless without achieving control of jetseparation.

[0056] These gases can be reinjected at a few points only into thedivergent body 4 of the nozzle to achieve jet separation according tothe invention.

[0057] The present invention can be adapted to the Vulcain 2 engine withminimal modifications. It is sufficient to modify the injection ring sothat it exhibits for example three injection points 5 instead of auniform distribution of injection. Furthermore, the divergent portion 4of the nozzle, which currently exhibits for the Vulcain 2 engine an arearatio R equal to 60 for a specific impulse of 433 seconds, could bereplaced with a divergent portion exhibiting an area ratio R of theorder of 140. The film-based cooling function could be replaced withradiative cooling, by virtue of a carbon/carbon nozzle extension knownper se.

[0058] For the Vulcain 2 engine, the value of Psep is of the order of0.22 bar and the recommended location for siting the points of injection5 is the cross section for which the pressure is equal to 0.4 bar. Thiscorresponds to a Mach number of 4 and an area ratio R of around 26. Thelocation is not very different from the current location of theinjection ring. The anticipated increase in the specific impulse is ofthe order of 12 seconds.

[0059] With altitude, the location of the source cross section wherenatural jet separation occurs migrates progressively downstream as theexternal pressure described decreases.

[0060] Furthermore, outside the atmosphere, secondary injection exhibitsno benefit, and even penalizes the performance of the engine.

[0061] A minimal procedure is to deactivate secondary injection when therocket leaves the atmosphere. In the Vulcain 2 engine, it is sufficientto add a valve which switches from a local injection mode to adistributed mode (film-based cooling).

[0062] Another solution is to arrange different injection points 5 whichare activated in succession in such a way as to optimize the operationof the nozzle at each instant. A solution of this type has already beenproposed, but for injection at continuous rings by the aforesaid patentU.S. Pat. No. 3,925,982.

[0063] The ignition of the engine constitutes a tricky phase duringwhich pressure transients are apt to exist. It is therefore desirable tominimize the considerable asymmetries of pressure which are apt to occurand which are apt to engender a high level of stress in the nozzle. Thetheoretically least favorable case is that for which the jet of thenozzle is momentarily entirely attached to the wall over one half of thenozzle and entirely separated over the other half thereof.

[0064] Owing to the violence of the unsteady fluctuations in the flowsduring start-up of a rocket engine, it is preferable, during thisstart-up, to use another embodiment which will be described hereinbelow.

[0065] Indeed, it is possible to solve the problem with injection of acontrol fluid during the very short start-up time of the engine, whichis of the order of a second. The point of impact of the injected fluidis close to the throat 3 of the nozzle, for example a distance of theorder of 0.1 D1 from the throat 3, D1 designating the length of the bodyof the nozzle 4, given that the pressure of the chamber is lower thanwhen the engine is at full throttle.

[0066] Finally, given that the jet is not organized, the stabilizing ofthe system requires a massive injection of fluid.

[0067] The device for stabilization on blast-off is represented in FIG.2. It can be used independently or otherwise of the jet separationdevice. It implements a plurality of injection tubes 10 parallel orotherwise to the axis of the nozzle and arranged downstream of thenozzle exit 8 and directed toward impact points 12. These tubes 10propagate fluid jets 11 in counter-current to the main stream, thepoints of impact 12 of these jets being situated slightly downstream ofthe throat 3 of the nozzle, for example a distance from the throat 3equal to 0.1 D1. These points of impact 12 distributed uniformly at oneand the same distance from the throat 3 of the nozzle produce a similareffect to that of the injection points 5, with the difference howeverthat the fluid, for example liquid nitrogen, which is projected createsa separation at each point of impact 12 through a mass entrainmenteffect along the boundary layer. The points of separation of the hotgases of the jet from the nozzle are stable by reason of the existenceof the impact points 12. It is advantageous to effect this injectionwith liquid nitrogen, since the counter-current injection rate may bevery high (for example 30 kg/s for each injection point for the Vulcain2 engine) during the short instant which proves to be necessary.Furthermore, the liquid nitrogen is transformed into gas when itencounters the hot gases originating from the combustion chamber 12,which means that the mass flux thus added artificially helps to reducethe phenomenon of spontaneous separation. Once full thrust has beenestablished, the nitrogen jet no longer penetrates into the body of thenozzle and it no longer has an influence on the operation of the engine.The stabilizing device is a ground-based device which is generallyarranged downstream of the exit 8 of the nozzle and which requires nomodification of the engine or of the launcher. It is apt to be used withnozzles which do or do not exhibit an injection device as represented inFIG. 1.

[0068] The invention is not limited to the exemplary embodimentsdescribed. In particular, jet separation could be initiated for exampleonward of a plurality of retractable inserts made of refractory materialintroduced radially into the wall of the nozzle.

1. A rocket engine nozzle comprising a system for controlling jet separation of the flow in the nozzle, wherein said control system exhibits a plurality of separation triggering elements (5, 10) arranged in such a way as to generate, from mutually spaced initiation points (9), distinct zones (6) of jet separation, so as to form a three-dimensional separation of the flow.
 2. The nozzle as claimed in claim 1, wherein the flow control system exhibits a device for injecting fluid through a wall of the nozzle, which exhibits, in at least one injection cross section substantially perpendicular to the axis of the nozzle, at least two independent injection orifices (5) distributed over the perimeter of the wall of the nozzle, each injection orifice (5) constituting a said separation triggering element inducing a said distinct zone (6) of jet separation.
 3. The nozzle as claimed in claim 2, wherein the injection orifices (5) are uniformly distributed over the perimeter of the wall of the nozzle (4).
 4. The nozzle as claimed in claim 3, wherein the injection orifices (5) are two in number and are diametrically opposed.
 5. The nozzle as claimed in claim 3, wherein the injection orifices (5) are 3 in number and are arranged at substantially 120° to one another over the perimeter of the nozzle (4).
 6. The nozzle as claimed in one of the preceding claims, wherein said injection cross section is arranged at a distance D from the throat (3) of the nozzle which is substantially less than the distance Do of spontaneous separation of the flow.
 7. The nozzle as claimed in claim 6, wherein the injection device exhibits a plurality of injectors (5) situated at different distances D, and a distributing device for feeding one or other of said injection cross sections (5), in such a way as to take into account the variation of said distance Do as a function of altitude.
 8. The nozzle as claimed in one of the preceding claims, wherein the flow control system exhibits an external stabilizing device integral with a ground-based installation and which exhibits, on the one hand, a number N(N≧2) of injection tubes (10) each of which constitutes a said separation triggering element, and which are distributed, preferably downstream of the nozzle (4), in such a way as to direct in counter-current to the main stream of the nozzle a stabilizing fluidic stream toward N impact points (12) situated downstream of the throat (3) of the nozzle (4), and on the other hand, a device (AL) for feeding the injection tubes (10) so as to feed them with fluid for a predetermined transient duration of ignition before takeoff, with a flow rate which is sufficient for each impact point (12) to induce a different zone of jet separation of the nozzle.
 9. The nozzle as claimed in claim 8, wherein the injection tubes (10) are parallel to the axis of the nozzle.
 10. The nozzle as claimed in either of claims 8 and 9, wherein the injection tubes (10) are arranged at the outlet of the nozzle (4) exit (8).
 11. The nozzle as claimed in one of claims 8 to 10, wherein the impact points (12) of the external stabilizing device are uniformly distributed over the perimeter of the wall of the nozzle.
 12. The nozzle as claimed in claim 11, wherein the impact points (12) of the external stabilizing device are two in number and are diametrically opposed.
 13. The nozzle as claimed in claim 11, wherein the impact points (12) of the external device are three in number and are arranged at substantially 120° to one another over the perimeter of the nozzle. 